Blade row, blade and turbomachine

ABSTRACT

A blade row for a turbomachine is disclosed. The blade row has an inner lateral wall and an outer lateral wall for bordering a hot gas channel in the turbomachine through which hot gas passes. At least one of the lateral walls has a rounded lateral wall front edge, which is provided with a circumferentially asymmetrical edge contouring that is elliptical or is comprised of one or more segments of a circle.

This application claims the priority of European Patent Application No.EP 13169058.8, filed May 24, 2013, the disclosure of which is expresslyincorporated by reference herein.

BACKGROUND AND SUMMARY OF THE INVENTION

The invention relates to a blade row of a turbomachine, a blade and aturbomachine.

To minimize secondary flows, blade rows in turbomachines, such as gasturbines and aircraft engines in particular, are often equipped with alateral wall contouring on the housing side and/or on the hub sidebordering a hot gas channel of the turbomachine through which a hot gasflows. A lateral wall front edge is traditionally designed withcircumferential symmetry in order to maintain an axial gap from anupstream blade row and therefore it is not uncontoured. U.S. PatentApplication Publication 2010/0172749 A1, however, describes a blade rowhaving a circumferentially asymmetrical lateral wall front edge,provided with wave-shaped edge contouring, which is comprised of avariety of rounded or adjusted surfaces in the direction of flow of thehot gas and/or the axial direction of the turbomachine in general.

The object of the invention is to create a blade row of a turbomachinethat has an alternative lateral wall front edge contouring. In addition,another object of the present invention is to create a blade for such ablade row and a turbomachine having such a blade row.

A blade row of a turbomachine according to the invention has an internallateral wall and an external lateral wall for bordering a hot gaschannel of the turbomachine through which a hot gas passes. At least oneof the lateral walls has a rounded lateral wall front edge, which isprovided with a circumferentially asymmetrical edge contouring, which iselliptical according to the invention or is comprised of one or moresegments of a circle.

The elliptical design of the edge contouring and/or the formation of thecircumferentially asymmetrical edge contouring from a plurality ofsegments of a circle, as seen in the direction of flow of the hot gasand/or in the axial direction of the turbomachine, causes a reduction insecondary flows in the area of the lateral wall front edge in comparisonwith a blade row having a circumferentially symmetrical lateral wallfront edge. In addition, the variation in a static pressure in thecircumferential direction in the hot gas channel and in the internalcooling channel is reduced due to the edge contouring on an internallateral wall front edge and/or a hub-side lateral wall front edge, thusleading to a reduction in local influx and efflux intensities, therebyreducing the mixing losses between the hot gas and the cool air.Furthermore, a leakage flow with its flow structures can be coordinatedwith a secondary flow in the blade row in a targeted manner.

In a preferred exemplary embodiment, the edge contouring is wave-shaped,as seen in the circumferential direction. The wave contour may be formedby two elevations and a recess situated between the elevations, forexample, such that the elevations extend over the lateral edges of theblades, considered respectively, so that two elevation sections areformed on each blade in its lateral edge area, each section togetherwith a corresponding elevation section on the neighboring blade formingone elevation. The elevations and recesses each relate to an idealcircumferentially symmetrical lateral wall front edge and yield afurther reduction in secondary flows.

An oncoming flow against the root side of the turbine blades of theblade row can be improved if root fillets of the turbine blades areinserted into the edge contouring.

The secondary flows can be reduced in the direction of flow over theentire lateral wall if the lateral wall is also contoured, and inparticular if the lateral wall contouring is then integrated into theedge contouring.

To reduce the structural complexity and manufacturing complexity, oneexemplary embodiment provides for the lateral wall to be designedwithout contour, such that the edge contouring tapers out into it.

To minimize a front axial gap despite the edge contouring, a frontlateral wall overhang is preferably free of contour and thus iscircumferentially symmetrical.

To reduce the structural complexity and manufacturing complexity, a rearlateral wall overhang may be free of contour and thus circumferentiallysymmetrical.

A blade according to the invention has a platform, the platform frontedge of which is rounded and has an edge contouring, which is ellipticalor is comprised of several segments of circles. A number of such bladesmake it possible to form a blade row of a turbomachine according to theinvention, which makes it possible to reduce secondary flows in the hotgas channel on the lateral wall end, static pressure differences andmixing losses. Furthermore, leakage flows can be passed through theblade row in a targeted manner.

A turbomachine according to the invention has at least one blade rowaccording to the invention. A turbomachine, such as a gas turbine and anaircraft engine in particular, is characterized by a superiorefficiency, because it permits reductions in lateral wall-side secondaryflows in the hot channel, static pressure differences and mixing losses.Furthermore, leakage flows can be passed through the blade row in atargeted manner. The at least one blade row is preferably situated onthe turbine end and in a low-pressure turbine of the turbomachine, forexample.

Preferred exemplary embodiments of the invention are explained ingreater detail below on the basis of schematic diagrams.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a perspective diagram of a partial region of a blade rowaccording to the invention with a first exemplary embodiment of a bladeaccording to the invention;

FIG. 2 shows a circumferential section of a blade row according to theinvention with a second exemplary embodiment of the blade according tothe invention;

FIG. 3 shows a side view of a third exemplary embodiment of the bladeaccording to the invention;

FIG. 4 shows a circumferential section of a blade row according to theinvention with a fourth exemplary embodiment of the blade according tothe invention;

FIG. 5 shows a perspective diagram of a fifth exemplary embodiment ofthe blade according to the invention; and

FIG. 6 shows a side view of the fifth exemplary embodiment of the bladeaccording to the invention from FIG. 5.

DETAILED DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a perspective diagram of a subsection of a blade row 1 of aturbomachine according to the invention, with a first exemplaryembodiment of the blades 2, 4 according to the invention. Theturbomachine is a gas turbine and is an aircraft engine in particular.The blade row 1 is preferably situated on the turbine end, for example,in a low-pressure turbine of the turbomachine. In the exemplaryembodiment shown here, the blades 2, 4 are rotor blades in particular,but they may also be guide blades, as illustrated in FIGS. 3 and 4 inparticular. The blades 2, 4 are turbine blades in particular.

The blade row 1 has a hub-side lateral wall 6 and a housing-side lateralwall, radially bordering a hot gas channel with a hot gas flowingthrough it essentially in the axial direction of the turbomachine. Forreasons of simplicity, only the hub-side and/or internal lateral wall 6is/are shown in FIG. 1 and the following FIGS. 2 to 6.

The hub-side lateral wall 6 is formed by internal and/or hub-sideplatforms 8, 10 of the blades 2, 4, which are numbered in FIG. 2 as anexample and are positioned side by side in the circumferential directionand from which turbine blades 12, 14 extend into the hot gas channel.The hub-side lateral wall 6 has a lateral wall front edge 16 and alateral wall rear edge 18, as seen in the direction of flow of the hotgas.

According to FIG. 1, the lateral wall front edge 16 forms a transitionalregion between the lateral wall 6 and a lateral wall overhang 20, whichis positioned in the front and/or upstream on the inside radially to thehub-side lateral wall 6. In the exemplary embodiment shown in FIG. 1,the lateral wall rear edge 18 of the lateral wall 6 forms a transitionalregion to a lateral wall overhang 22, which is positioned at the rearand/or downstream, on the inside radially to the front lateral walloverhang 20.

As shown in FIG. 2, according to the formation of the lateral wall 6,the lateral wall front edge 16, the lateral wall rear edge 18, the frontlateral wall overhang 20 and the rear lateral wall overhang 22 areformed by individual platform front edges 24, 26, platform rear edges28, 30, front platform overhangs 32, 34 and/or rear platform overhangs36, 37 positioned side by side.

As shown in FIG. 1, the lateral wall front edge 16 is designed to berounded and is provided with edge contouring. The edge contouring issuch that the lateral wall front edge 16 is circumferentiallyasymmetrical. The edge contouring is preferably uniformly wave-shaped,based on an ideal, circumferentially symmetrical lateral wall frontedge, wherein an elevation 38, 40 extending onto both platforms 8, 10and a recess 42 provided between the elevations 38, 40 are both arrangedin the lateral edge area of the blades 2, 4. Thus two correspondingelevation sections 44, 46, which, taken separately and isolated from thelateral wall front edge 16, form the elevations 38, 40 are situated ineach lateral edge region of the blades 2, 4. The turbine blades 12, 14are guided into one of the elevation sections 46, each with an oncomingflow edge region 48 of their rounded root in the exemplary embodimentshown here.

The recess 42 is formed between the elevation sections 44, 46 and ispreferably situated centrally with the respective platform front edge24, 26 (see FIG. 2). The recess 42 is offset radially toward the insidein comparison with an ideal lateral wall edge line and/or hot gaschannel line, so that in its lowest section, it comes in contact with afoot 50 on the outside radially of the front lateral wall overhang 20.The front lateral wall overhang 20, which has circumferential symmetry,thus develops directly into the recess 42 with its outer foot 50 in atleast some sections. However, the elevations 38, 40 and/or elevationsections 44, 46 are at a distance from the outer foot 50, such that freeareas 52, 54 extending radially are not contoured between the outer foot50 and the elevations 38, 40 and/or elevation sections 44, 46.

In the axial direction of the turbomachine, and thus in simplifiedterms, in the direction of flow and/or as seen from the front lateralwall overhang 20 to the lateral wall 6, the edge contouring iselliptical or is comprised of a plurality of segments of circles withdifferent radii.

The edge contouring develops into a circumferentially asymmetricallateral wall 6 and/or a circumferentially asymmetrical lateral wallcontouring. In the exemplary embodiment shown in FIG. 1, the lateralwall contouring has two peaks, which extend over the lateral edge areaof the blades 2, 4 to the neighboring platform 8, 10. These peaks arearranged one after the other in succession, and each is shown only as apeak section 56, 58 due to the perspective. The front peak is at adistance from a suction-side area 60 of the root fillet by way of achannel, which has not been assigned a number. The rear peak extendsfrom an uncontoured lateral wall section 62.

The lateral wall contouring tapers out in the direction of the lateralwall rear edge 18. The lateral wall rear edge 18 is not contoured andtherefore has circumferential symmetry. Likewise the rear lateral walloverhang 22 is not contoured and therefore has circumferential symmetry.

FIG. 2 shows a hub-side circumferential section of a turbine blade 1,which is formed from a plurality of blades 2, 4 according to theinvention, such as turbine rotor blades according to a second exemplaryembodiment. The blade row 1 has a circumferentially symmetrical frontlateral wall overhang 20, a circumferentially asymmetrical lateral wallfront edge 16, a circumferentially asymmetrical lateral wall 6, acircumferentially symmetrical lateral wall rear edge 18 and acircumferentially symmetrical rear lateral wall overhang 22.

The circumferentially asymmetrical lateral wall front edge 16 is roundedand has a wave-shaped edge contouring in the circumferential direction,which appears elliptical in the direction of flow and/or as seen fromthe front lateral wall overhang 20 to the lateral wall 6 or is comprisedof several segments of circles with different radii. As in the exemplaryembodiment according to FIG. 1, the edge contouring develops into acircumferentially asymmetrical lateral wall contouring of the lateralwall 6.

The lateral wall contouring of the lateral wall 6 according to theexemplary embodiment shown in FIG. 2 has two peaks 64, 66 near a suctionside 67 of each turbine blade 12, 14. The peaks 64, 66 are spaced adistance apart from one another in the direction of flow and eachextends beyond the lateral edge regions of platforms 8, 10 of theneighboring blades 2, 4. A region of the lateral wall 6 and/or of theplatforms 8, 10 not provided with peaks 64, 66 is provided with a valley68 extending beyond the lateral edge regions. Thus each platform has twosuction-side peak sections 70, 72 and two more-or-less pressure-sidepeak sections 74, 76, which together with the corresponding peaksections of the neighboring blades 2, 4, form the peaks 64, 66. Thepeaks 64, 66 are each spaced a distance apart from a suction-side region60 of a root fillet through a channel. Furthermore, the root fillet isguided onto the respective neighboring platform 8, 10 on the suctionside.

FIG. 3 shows a section through an exemplary embodiment of the blade 2according to the invention, which is designed as a guide blade, such asa turbine guide blade of a turbomachine. In significant contrast to theexemplary embodiments illustrated in FIGS. 1 and 2, this blade 2 doesnot have any front lateral wall overhang or any rear lateral walloverhang.

However, the blade 2 shown in FIG. 3, like the preceding exemplaryembodiments, does have a hub-side circumferentially asymmetricalplatform front edge 24, a hub-side circumferentially asymmetricalplatform 8 and a hub-side circumferentially symmetrical platform rearedge 28.

The circumferentially asymmetrical platform front edge 24 is rounded andhas a wave-shaped edge contouring, as seen in the circumferentialdirection, which appears elliptical, as seen in the direction of flow,or is comprised of multiple segments of circles with different radii. Asin the exemplary embodiments according to FIGS. 1 and 2, the edgecontouring develops into the platform 8 and/or a platform contouring.

With its platform contouring, the platform 8 forms a partial region of acircumferentially asymmetrical lateral wall contouring. Starting from anideal hot gas channel line 78, the platform contouring is such that thelateral wall contouring has a front peak 64 and a rear peak 66 betweentwo turbine blades 12, the rear peak being designed to be lower incomparison with the front peak 64 and being connected to the front peak64 by means of a section 80 with a reduced height in relation to thepeaks 64, 66. The platform contouring tapers out in the direction of thecircumferentially symmetrical platform rear edge 28, which is preferablyangular in the exemplary embodiment shown here.

FIG. 4 shows a hub-side circumferential section of a blade row 1 formedfrom a plurality of the blades 2, 4 according to the invention, such asturbine guide blades according to a fourth exemplary embodiment.According to the preferred embodiment of the blades 2, 4 as turbineguide blades, the blade row 1 does not have a front lateral walloverhang or a rear lateral wall overhang, but it does have acircumferentially asymmetrical lateral wall front edge 16, acircumferentially asymmetrical lateral wall 6 and a circumferentiallysymmetrical lateral wall rear edge 18.

The circumferentially asymmetrical lateral wall front edge 16 is roundedand has a wave-shaped edge contouring as seen in the circumferentialdirection, appearing elliptical, as seen in the direction of flow and/orfrom the front lateral wall overhang to the lateral wall, or iscomprised of multiple segments of circles with different radii. In theexemplary embodiments according to the preceding FIGS. 1 to 3, the edgecontouring develops into the circumferentially asymmetrical lateral wall6 and/or its lateral wall contouring.

The lateral wall contouring has a peak 64 which extends between rootfillets 82, 84 of the respective neighboring turbine blades 12, 14 anddevelops into the root fillets 82, 84 at the height of an ideal hot gaschannel line.

In the area of the suction-side root fillet 84, the peak 64 extendsaround an uncontoured lateral wall section 88, from which thesuction-side root fillet 84 emerges smoothly. Furthermore, the lateralwall contouring has a front valley 68 and a rear valley 86, each ofwhich is between two turbine blades 12, 14, separated from one anotherby the peak 64.

As also shown in FIG. 4, the lateral wall contouring tapers out from theoutflow edges 90 of the turbine blades 12, 14 in an approximate axialposition, so that a noncontoured lateral wall section 62 is formedbetween the outflow edges 90 and the lateral wall rear edge 18.

FIGS. 5 and 6 show an exemplary embodiment of the blade 2 according tothe invention, which is designed as a rotor blade, such as a turbinerotor blade of a turbomachine. The blade 2 has a hub-side front platformoverhang 32, a hub-side platform front edge 24, a hub-side platform 8, ahub-side platform rear edge 28 and a hub-side rear platform overhang 36.

In substantial contrast with the exemplary embodiments according toFIGS. 1 to 4, the platform 8 is not contoured but instead is designed tobe circumferentially symmetrical.

In further substantial contrast, the rear platform overhang 36 continuesan ideal hot gas channel line 78 of the platform 8, so that in contrastwith the exemplary embodiments according to FIGS. 1, 2 and 3, it is notoffset radially toward the inside in comparison with the platform 8 butinstead it forms the platform rear edge 28. In the exemplary embodimentsaccording to FIGS. 1, 2 and 3, the front platform overhang 32, the rearplatform overhang 36 and the platform rear edge 28 are allcircumferentially symmetrical.

The platform front edge 24 is rounded and has a circumferentiallyasymmetrical edge contouring, as seen in the circumferential direction,appearing elliptical in the direction of flow and/or as seen from thefront platform overhang 32 to the platform 8, or is comprised of aplurality of segments of circles with different radii. As in thepreceding exemplary embodiments, the edge contouring may be wave-shapedin the circumferential direction. The edge contouring develops into thecircumferentially symmetrical platform contouring and/or tapers out intothat.

It should be pointed out that although FIGS. 1 to 6 are based only onthe hub side and/or inner lateral wall 6/platforms 8, 10, anyconsiderations and/or contouring and/or noncontourings may of coursealso be applied in the case of an outer and/or housing-side lateralwall/platform. In particular, both the hub-side lateral wall 6 and thehousing-side lateral wall as well as the blade platforms 8, 10 requiredfor this purpose may be contoured accordingly for bordering a hot gaschannel of a turbomachine.

A blade row for a turbomachine is disclosed, having an inner lateralwall and an outer lateral wall for bordering a hot gas channel of theturbomachine through which hot gas flows, wherein at least one of thelateral walls has a rounded lateral wall front edge, which is providedwith a circumferentially asymmetrical edge contouring that is ellipticalor is comprised of a plurality of segments of circles, a blade for sucha blade row and a turbomachine.

LIST OF REFERENCE CHARACTERS

-   1 Blade row-   2 Blade-   4 Blade-   6 Inner/hub-side lateral wall-   8 Platform/blade platform-   10 Platform/blade platform-   12 Turbine blade-   14 Turbine blade-   16 Lateral wall front edge-   18 Lateral wall rear edge-   20 Front lateral wall overhang-   22 Rear lateral wall overhang-   24 Platform front edge-   26 Platform front edge-   28 Platform rear edge-   30 Platform rear edge-   32 Front platform overhang-   34 Front platform overhang-   36 Rear platform overhang-   37 Rear platform overhang-   38 Elevation-   40 Elevation-   42 Recess-   44 Elevation section-   46 Elevation section-   48 Region of a root fillet on the side of the oncoming flow edge-   50 Outer foot-   52 Free area-   54 Free area-   56 Peak section-   58 Peak section-   60 Suction-side region of a root fillet-   62 Noncontoured lateral wall section-   64 Peak-   66 Peak-   67 Suction side-   68 Valley-   70 Peak section-   72 Peak section-   74 Peak section-   76 Peak section-   78 Ideal hot gas channel line/ideal platform line-   80 Section with a reduced height-   82 Root fillet-   84 Root fillet-   86 Valley-   88 Noncontoured lateral wall section-   90 Outflow edge

As also discussed above, the foregoing disclosure has been set forthmerely to illustrate the invention and is not intended to be limiting.Since modifications of the disclosed embodiments incorporating thespirit and substance of the invention may occur to persons skilled inthe art, the invention should be construed to include everything withinthe scope of the appended claims and equivalents thereof.

What is claimed is:
 1. A blade row of a turbomachine, comprising: aninner lateral wall; an outer lateral wall; wherein the inner lateralwall and the outer lateral wall border a hot gas channel of theturbomachine through which hot gas flows; and a plurality of turbineblades disposed between the inner lateral wall and the outer lateralwall; wherein at least one of the inner lateral wall and the outerlateral wall has a rounded lateral wall front edge which includes acircumferentially asymmetrical edge contouring; and wherein the edgecontouring is elliptical or is comprised of one or more segments of acircle.
 2. The blade row according to claim 1, wherein the edgecontouring is wave-shaped in a circumferential direction.
 3. The bladerow according to claim 1, wherein respective root fillets of theplurality of turbine blades are disposed in the edge contouring.
 4. Theblade row according to claim 1, wherein the edge contouring isintegrated into a lateral wall contouring of the at least one of theinner lateral wall and the outer lateral wall.
 5. The blade rowaccording to claim 1, wherein the at least one of the inner lateral walland the outer lateral wall is free of a contour and the edge contouringtapers into the at least one of the inner lateral wall and the outerlateral wall.
 6. The blade row according to claim 1, wherein a lateralwall rear edge of the at least one of the inner lateral wall and theouter lateral wall is free of a contour.
 7. The blade row according toclaim 1, wherein a front lateral wall overhang of the at least one ofthe inner lateral wall and the outer lateral wall is free of a contour.8. The blade row according to claim 1, wherein a rear lateral walloverhang of the at least one of the inner lateral wall and the outerlateral wall is free of a contour.
 9. A turbomachine having a blade rowaccording to claim 1.